Gas turbine shroud structure

ABSTRACT

There is disclosed a gas turbine shroud structure comprising: a shroud support component  12  attached to an inner surface of a gas turbine casing  3 ; a shroud segment  14  divided in a peripheral direction and supported by the inner surface of the shroud support component; and a heat-resistant restricting spring  18  held between the shroud segment and shroud support component to urge the shroud segment inwards in a radial direction. The shroud segment  14  is formed of a ceramic composite material, and includes a coating layer  15  having heat insulation and impact absorption effect on an inner surface of the segment.

BACKGROUND OF THE INVENTION

1. Technical Field of the Invention

The present invention relates to a shroud structure which shrouds aturbine rotator blade of a gas turbine.

2. Description of the Related Art

In a gas turbine, an air compressor compresses air, a combustor combustsfuel in the compressed air, combustion gas drives the turbine, and theair compressor is driven by the driving force.

When an operation temperature of the gas turbine is raised, efficiencyis improved. Therefore, the turbine rotator blade, and a turbine shroudand turbine casing which shroud the blade is cooled, and the gas turbineis operated at a high temperature, so that the efficiency has heretoforebeen improved.

In recent years, from a necessity of enhancement of capability requiredfor development of a jet-engine and consideration to environment, and ithas been an important problem to reduce a cooling air amount as well asan engine weight.

A heat-resistant metal has heretofore been used in the turbine shroud,but in recent years a material particularly superior in ahigh-temperature characteristic has been applied with the hightemperature of a turbine portion. Moreover, a thermal barrier coatingwhich is a thermal interference layer is sometimes applied. However, forthe metal turbine shroud, it has already been impossible to furtherreduce the cooling air or weight. For the present situation in which thetemperature tends to be further raised, the amount of the cooling airhas to be increased. Moreover, the weight cannot be reduced. In otherwords, metal components have already reached their limitations, andfurther reduction of the cooling air amount or weight cannot be expectedin the situation.

To solve these problems, a shroud of a ceramic composite material (CMC)has also been proposed (e.g., Japanese Patent Application Laid-Open Nos.10-103013, 10-103014).

In “Gas Turbine Shroud Structure” of the Japanese Patent ApplicationLaid-Open No. 10-103013, as shown in FIG. 1, a cylindrical shroud 1 isconstituted of a plurality of segments 2 which are divided a peripheraldirection to have circular-arc plate shapes. Opposite edges of eachsegment along a main-stream gas direction are supported by a supportmember 4 fixed on an inner peripheral side of a gas turbine casing 3,and each segment 2 is constituted of ceramic and has a double platestructure which includes an outer peripheral side plate portion 2 a andinner peripheral side plate portion 2 b. It is to be noted that in thisdiagram, reference numeral 5 denotes a first-stage stationary blade, and6 denotes a first-stage rotator blade.

However, in the shroud in which such two-dimensional plate-shaped CMCmaterial is used, delamination occurs, a high stress is generated in aportion formed by combining two components 2 a, 2 b by deformation at ahigh temperature, and there is a problem of high possibility that theportion is destroyed.

On the other hand, in “Gas Turbine shroud Structure” (FIG. 2) of theJapanese Patent Application Laid-Open No. 10-103014, the respectivesegments 2 in FIG. 1 are constituted of ceramic, and a section of thesegment has a hollow constitution which has a square, trapezoidal, I oranother shape by braiding.

However, in the shroud using the braided CMC material, many fiberscannot be included in the braided material. When the segment islengthened, thermal stress increases. There is a problem that it isdifficult to establish strength.

Moreover, in each of the above-described prior-art CMC shrouds,a-countermeasure against mismatch by a thermal expansion difference isnot considered for any connected portion to the metal component. Thereis also a high possibility that a restricted portion is destroyed.Moreover, since a sealing property with the metal component is notconsidered, the cooling air leaks and deteriorates the capability.Furthermore, a shroud main body is directly exposed to a main-streamgas, and it is therefore difficult to use the shroud main body inhigh-temperature environment in which a main-stream gas temperatureexceeds 1200° C. for a long time. Moreover, there has been a possibilitythat the whole shroud is destroyed at a collision time with a rotatorblade tip end.

SUMMARY OF THE INVENTION

The present invention has been developed to solve the above-describedvarious problems. That is, an object of the present invention is toprovide a gas turbine shroud structure in which a shroud main body isnot directly exposed to a main-stream gas exceeds 1200° C. and which canavoid destruction of a shroud even in collision of a rotator blade tipend, handle a thermal expansion difference, prevent generation of a highthermal stress, and enhance gas turbine capability.

According to the present invention, there is provided a gas turbineshroud structure comprising: a shroud support component (12) attached toan inner surface of a gas turbine casing (3); a shroud segment (14)divided in a peripheral direction and supported by the inner surface ofthe shroud support component; and a heat-resistant restricting spring(18) held between the shroud segment and shroud support component tourge the shroud segment inwards in a radial direction, wherein theshroud segment (14) is formed of a ceramic composite material, andincludes a coating layer (15) having heat insulation and impactabsorption effect on an inner surface of the segment.

According to a preferred embodiment of the present invention, in theceramic composite material, a laminated and stitched two-dimensionalfabric, or a fabric in which the fiber is three-dimensionally orientedis used.

According to the constitution of the present invention, the shroudsegment (14) is formed of the ceramic composite material using thelaminated and stitched two-dimensional fabric or the fabric in which thefiber is three-dimensionally oriented, and therefore delamination can beprevented. Moreover, since the coating layer (15) having the heatinsulation and impact absorption effect is formed on the inner surfaceof the shroud segment, the shroud main body is not directly exposed tothe main-stream gas, and can be used in a high-temperature environmentat a main-stream gas temperature exceeding 1200° C. for a long time, anddestruction of a shroud can be avoided even in collision with a rotatorblade tip end.

Moreover, since the heat-resistant restricting spring (18) is heldbetween the shroud segment (14) and shroud support component (12) tourge the shroud segment inwards in a radial direction, the thermalexpansion difference can be handled, and high thermal stress can beprevented from being generated. Furthermore, a gap by backlash of theshroud support component and shroud segment is suppressed, and the leakamount of the cooling air can be reduced, so that the capability isimproved.

According to the preferred embodiment of the present invention, therestricting spring (18) binds the shroud segment (14) to the shroudsupport component (12) and has elasticity which can follow the thermalexpansion difference from the shroud support component in the radialdirection.

According to the constitution, a high thermal stress is prevented frombeing generated by the thermal expansion difference of the shroudsegment (14) in axial and radial directions, the gap by the clearance inthe radial direction can be prevented by the restricting spring (18),and the leak of the cooling air can be suppressed.

Moreover, the coating layer (15) is ceramic including pores, or ceramiccomposite material. The ceramic including the pores or ceramic compositematerial is preferably formed by ceramic containing mullite, SiC, Al₂O₃,zircon, SiO₂, ZrC, HfC which are main components.

According to the constitution, the coating layer can be formed to havethe heat insulation and impact absorption effect.

Other objects and advantageous characteristics of the present inventionwill be apparent from the following description with reference to theaccompanying drawings.

BRIEF DESCRIPTION OF DRAWINGS

The patent or application file contains at least one drawing executed incolor. Copies of this patent or patent application publication withcolor drawing(s) will be provided by the Office upon request and paymentof the necessary fee.

FIG. 1 is a diagram showing a related-art gas turbine shroud structure;

FIG. 2 is a diagram showing another segment of the related-art gasturbine shroud structure;

FIGS. 3A to 3D are sectional views showing the gas turbine shroudstructure according to the present invention;

FIGS. 4A and 4B are basic concept diagrams of the gas turbine shroudstructure according to the present invention;

FIG. 5 is a temperature distribution diagram of the gas turbine shroudstructure according to the present invention;

FIG. 6 is a perspective view of the shroud segment of the presentinvention;

FIG. 7 shows an embodiment showing an adhesion strength of the coatinglayer of the present invention; and

FIGS. 8A and 8B show photographs before and after a heat cycle test ofthe coating layer of the present invention.

DESCRIPTION OF THE PREFERRED EMBODIMENT

A preferred embodiment of the present invention will be describedhereinafter with reference to the drawings. It is to be noted thatcommon portions in the respective drawings are denoted with the samereference numerals, and redundant description is omitted.

FIGS. 3A to 3D are sectional views showing the gas turbine shroudstructure according to the present invention. As shown in the drawings,the gas turbine shroud structure of the present invention includes ashroud support component 12, shroud segment 14, and restricting spring18.

The shroud support component 12 is a metal component, and is attached tothe inner surface of a gas turbine casing 3 via bolts.

The shroud segment 14 is a circular arc member divided in a peripheraldirection, and is fixed to the inner surface of the shroud supportcomponent 12. As shown in FIGS. 3B, 3C, 3D, the shroud segment 14 isformed of a ceramic composite material using a laminated and stitchedtwo-dimensional fabric, or a fabric in which the fiber isthree-dimensionally oriented.

FIG. 3B shows an orthogonal three-dimensional fabric which includesfibers in a plate thickness direction (diametric direction of an engine)and which is reinforced by two-dimensional stitching in the platethickness direction for a purpose of reinforcing a destruction mode of ashroud thickness direction (diametric direction of the engine) in whicha hook portion is assumed.

FIGS. 3C and 3D show the fabric in which the fiber constituting the hookportion is continuously bent and disposed to reinforce a shroudthickness direction and the continued fiber reinforces the shroudthickness direction (diametric direction of the engine).

Moreover, the shroud segment 14 includes a coating layer 15 having heatinsulation and impact absorption effect on the inner surface thereof.The coating layer 15 is ceramic including pores, or ceramic compositematerial, and is preferably formed by the ceramic containing mullite,SiC, Al₂O₃, zircon, SiO₂, ZrC, HfC which are main components.

As a coating method, ceramic molding methods such as spray coating andslurry methods, and methods for matrix molding of CMC, such as CVI andPIP can be used.

The restricting spring 18 is formed of a heat-resistant thin metal, andis held between the shroud segment 14 and shroud support component 12 tourge the shroud segment 14 inwards in a radial direction. Thisrestricting spring 18 has elasticity which can follow a thermalexpansion difference of the shroud segment 14 in the radial direction.

According to the above-described constitution of the present invention,the shroud segment 14 is formed of the ceramic composite material usingthe laminated and stitched two-dimensional fabric, or the fabric inwhich the fiber is three-dimensionally oriented, so that delaminationcan be prevented. Moreover, since the coating layer 15 having the heatinsulation and impact absorption effect is formed on the inner surfaceof the shroud segment, the shroud main body is not directly exposed tomain-stream gas. The shroud can be used in a high-temperatureenvironment at a main-stream gas temperature exceeding 1200° C. for along time, and destruction of the shroud can be avoided even incollision with a rotator blade tip end.

Moreover, the heat-resistant restricting spring 18 is held between theshroud segment 14 and shroud support component 12 to urge the shroudsegment inwards in the radial direction. Therefore, the thermalexpansion difference can be handled, a high thermal stress can beprevented from being generated, a gap by the clearance between theshroud support component and shroud segment is suppressed, and a leakamount of cooling air can be reduced, so that capability is enhanced.

Moreover, according to this constitution, the high thermal stress can beprevented from being generated by the thermal expansion difference ofthe shroud segment 14 in axial and radial directions, the gap by thebacklash in the radial direction is suppressed by the restricting spring18, and the leak amount of the cooling air can be reduced, so that thecapability is enhanced.

The coating layer 15 is formed of the ceramic including the pores or theceramic composite material, and can be the coating layer which has theheat insulation and impact absorption effect.

The preferred embodiment of the present invention will be describedhereinafter in detail.

1. An aerial demand is expected to be twice the existing demand around2010, and a CO₂ discharge amount unavoidably becomes twice the existingamount with an extension of the existing technique. Therefore, to reducean influence on the environment, the CO₂ discharge amount needs to bereduced to the existing level, and there has been a demand for rapidenhancement of an engine associated technique and airplane bodyassociated technique.

A ceramic matrix composite (CMC) is lightweight and superior in heatresistance. Therefore, when the material is applied to a jet enginecomponent, weight reduction of a jet engine and reduction of a fuelconsumption ratio can be anticipated. Many researches for application toaerospace components of CMC have been performed, and works haveintensively been performed for practical use in recent years.

An object of the present invention is to apply CMC having about ¼specific gravity as compared with a nickel group alloy and superior inheat resistance to a turbine shroud component which is directly exposedto a main-stream high-temperature gas in the jet engine for ansupersonic airplane, and to realize the CO₂ discharge amount reductionby a technical development concerning the reduction of the engine weightand fuel consumption ratio. When the CMC is applied to the turbineshroud, there are the following technical problems.

A manufacturing process which keeps a material strength while satisfyinga three-dimensional shape

A fastening structure of the CMC and metal components which have a largethermal expansion difference

Durability in a gas at a high temperature exceeding 1500° C.

In the embodiment of the present invention, to evaluate manufacturingprocess required for applying the CMC to the turbine shroud andcharacteristics for forming the coating layer, a shroud model is made ontrial, and adequacy of a manufacturing process is confirmed. Moreover, atest specimen in the form of a CMC flat plate on which the coating layeris formed is made on trial, and adhesion and heat resistance cyclecharacteristics are evaluated.

2. Study of CMC Shroud Structure

A basic concept of the CMC shroud of the present invention is shown inFIGS. 4A and 4B. First, two-dimensional and three-dimensional fabricsare considered as a fabric structure of the CMC shroud model. However,since interlayer strength is a problem in a two-dimensional laminatematerial, a flat plate fabric having an orthogonal three-dimensionalfiber orientation is used. For the flat plate fabric, a method ofmanufacturing a rectangular parallelepiped fabric and subsequentlybending the fabric in a circular arc shape was selected. For the fiberfor use in the flat plate fabric, Tyranno ZMI grade having a fiberdiameter of 11 μm and manufactured by Ube Industries, Ltd. was used.

After weaving the ZMI fiber into the flat plate fabric having theorthogonal three-dimensional fiber orientation, the surface of the fiberwas coated with carbon in about 0.1 to 0.2 μm by CVD method. For amatrix, the matrix of SiC was formed by a hybrid process constituted bycombining a chemical vapor infiltration (CVI) method and polymerimpregnation and pyrolysis (PIP) method.

The existing heat resistance of the CMC is in a range of 1200 to 1400°C., and the CMC needs to be protected from a main-stream gas at a hightemperature exceeding 1500° C. Therefore, the CMC shroud is coated asshown in FIG. 4B. For the coating, a mullite based material having athermal expansion close to that of the CMC was selected.

For the shape of the CMC shroud and the coating thickness, validity wasstudied and set by FEM analysis. A temperature distribution in a 1 mmcoating layer is shown in FIG. 5. The shroud is protected from thehigh-temperature gas by the coating layer, and a shield effect of heatat about 150° C. is also obtained. The CMC shroud made on trial usingthe above-described process is shown in FIG. 6. A trial manufacturingresult of the CMC shroud was satisfactory, and the adequacy of themanufacturing process in the above-described process was confirmed.

3. Coating Layer Evaluation Test Method

To evaluate the adhesion and heat resistance cycle characteristic in thecoating of the CMC, a CMC test specimen in the form of the flat platehaving the orthogonal three-dimensional fiber orientation was produced.A dimension of each test specimen is shown as follows.

Bonding strength test specimen shape: length of 20 mm, width of 20 mm,thickness of CMC portion of 4 mm, coating layer of 1 mm

Thermal cycle test specimen shape: length of 50 mm, width of 50 mm,thickness of CMC portion of 4 mm, coating layer of 1 mm

The mullite based material forming the coating layer of each testspecimen was disposed by spray coating.

In bonding strength test, metal rods were bonded on coating and CMCsides so as to hold the test specimen, and attached to a tensile testerto conduct the test.

Moreover, in the thermal cycle test, the test specimen was held by apin, and fixed to a cooling holder. The temperature of the back surfaceof the test specimen was monitored by a thermocouple, and thetemperature of a heating surface was monitored by a pyrometer, while thetest specimen was heated by a burner to conduct the test. For thetemperature, a temperature state assumed at a study time of thestructure of FIG. 5 was simulated, the coating surface (heating surface)temperature was controlled at 1350° C., and the temperature of a coolingsurface on the CMC side (back surface) was controlled at about 900° C.

4. Test Result and Consideration

FIG. 7 shows bonding strength test result. An bonding strength is shownin comparison with that of the existing coating on the metal shroudcomponent. It was confirmed that a satisfactory adhesion was obtained ascompared with the existing coating.

Conditions of the test specimen after 500 heat cycles are shown in FIGS.8A and 8B. In the beginning, a very small number of cracks capable ofbeing confirmed by a magnifying glass were generated in a boundarysurface of the coating layer and CMC after several cycles. However, evenafter 500 cycles, the cracks did not proceed much, and the debonding didnot occur. This was supposedly because micro cracks were generated inthe coating and the thermal stress in the coating was relaxed.

5. The manufacturing process required for applying the CMC to theturbine shroud and the CMC on which the coating layer was formed wereevaluated. As a result, the CMC shroud model was manufactured on trial,the manufacturing process property from the CMC using the flat platefabric having the orthogonal three-dimensional fiber orientation wasconfirmed, and the adequacy of the applied manufacturing process wasconfirmed.

As described above, in the present invention, the ceramic compositematerial is used, including a structure into which a heat interferencelayer at the rubbing time with the rotator blade tip end is introducedby the coating layer and which protects the CMC functioning as astructure member. In the ceramic composite material, the two-dimensionalfabric is laminated and stitched in the CMC portion which is thestructure member, or the fabric in which the fiber isthree-dimensionally oriented is used. Moreover, during the fasteningwith the metal component, the material is bound to the metal componenthaving a thermal elongation difference via the spring component.

Since the coating layer is introduced, the material having low thermalconductivity and including the pores is constituted. Thereby, the heatshield effect is obtained. Moreover, the interference layer at therotator blade tip end rubbing time is constituted, and the CMCfunctioning as the structure member can be protected from a thermalstress and structure load.

Moreover, for the CMC portion as the structure member, the laminated andstitched two-dimensional fabric, or the fabric in which the fiber isthree-dimensionally oriented is used. Thereby, the delamination which isa problem in a two-dimensional material or braiding material can beavoided.

Furthermore, for the fastened portion to the metal component, thecomponent is bound to the metal component having a thermal elongationdifference via the spring component, so that mismatch by a thermaldeformation difference can be eliminated and destruction can be avoided.

Therefore, the gas turbine shroud structure of the present invention haseffects that the delamination can inherently be prevented and the shroudmain body is not directly exposed to the main-stream gas. The structurecan be used in the high-temperature environment in which the main-streamgas temperature exceeds 1200° C. for a long time. Even when the rotatorblade tip end rubs with the shroud, the destruction of the shroud can beavoided. The thermal expansion difference can be handled, and the highthermal stress can be prevented from being generated.

It is to be noted that some preferred embodiments of the presentinvention have been described, but it could be understood that the scopeof rights involved in the present invention is not limited to theseembodiments. Contrarily, the rights scope of the present inventionincludes all improvements, modifications, and equivalents included inthe attached claims.

1. A gas turbine shroud structure comprising: a shroud support componentattached to an inner surface of a gas turbine casing; a shroud segmentdivided in a peripheral direction and supported by the inner surface ofthe shroud support component; and a heat-resistant restricting springheld between the shroud segment and the shroud support component to urgethe shroud segment inwards in a radial direction, wherein the shroudsegment is formed of a ceramic composite material, and includes acoating layer having heat insulation and impact absorption effect on aninner surface of the segment, wherein the ceramic composite materialcomprises laminated and stitched two-dimensional fabric, or fabric inwhich a fiber is three-dimensionally oriented.
 2. The gas turbine shroudstructure according to claim 1, wherein the restricting spring binds theshroud segment onto the shroud support component and has an elasticitywhich can follow a thermal expansion difference from the shroud supportcomponent in a radial direction.
 3. The gas turbine shroud structureaccording to claim 1, wherein the fabric in which the fiber isthree-dimensionally oriented is an orthogonal three-dimensional fabricin which the fiber is inserted in a plate thickness direction, or afabric in which the fiber constituting a hook portion is bent andoriented to continuously reinforce a shroud thickness direction and theshroud thickness direction is reinforced by the continued fiber.
 4. Thegas turbine shroud structure according to claim 3, wherein the coatinglayer is a ceramic including pores or a ceramic composite material,wherein each of the ceramic including pores and the ceramic compositematerial is formed by ceramic containing one or more main componentsselected from the group consisting of mullite, SiC, Al₂O₃, zircon, SiO₂,ZrC, and HfC.
 5. The gas turbine shroud structure according to claim 1,wherein the laminated and stitched two-dimensional fabric, and thefabric in which a fiber is three-dimensionally oriented, are rectangularparallelepiped fabric subsequently bent in a circular arc shape.
 6. Agas turbine shroud structure comprising: a shroud support componentattached to an inner surface of a gas turbine casing; a shroud segmentdivided in a peripheral direction and supported by the inner surface ofthe shroud support component; and a heat-resistant restricting springheld between the shroud segment and the shroud support component to urgethe shroud segment inwards in a radial direction, wherein the shroudsegment is formed of a ceramic composite material, and includes acoating layer having heat insulation and impact absorption effect on aninner surface of the segment, wherein the ceramic composite materialcomprises laminated and stitched two-dimensional fabric, or fabric inwhich a fiber is three-dimensionally oriented, wherein the coating layeris a ceramic including pores or a ceramic composite material, whereineach of the ceramic including pores and the ceramic composite materialis formed by ceramic containing one or more main components selectedfrom the group consisting of mullite, SiC, Al₂O₃, zircon, SiO₂, ZrC, andHfC.